Supersonic aircraft shock wave energy recovery system

ABSTRACT

This invention outlines excitation means to transform the linear momentum of an underwing energized jet into rotational form in a selective manner to provide an asymmetric shear layer to increase compression wave reflection from the forward undersurface of a supersonic wing. The wing compression energy is thereby recovered into useful work as an increase in pressure on the upward reflexed wing backside. The upper surface of the shear layer is comprised of an array of vortices whose rotation is opposite to the wing circulation, providing the required angular momentum reaction. The upper wing surface is flat to avoid generation of waves and an adverse angular momentum reaction above the wing. The vortices below the wing are compressed by the underwing pressure, comprising a pressure shield to enhance the reflection. The shear layer/vortex array grows in the stream direction due to augmented mixing with the underwing gap flow, which is turned and deflected upwards to provide a further increase in pressure on the upwards reflexed wing backside. Fuselage bow shock energy is also recovered into useful work by a forward ring reflecting the conical shock inwards onto a suitably inclined shoulder. An extendable nose spike allows the ring to intercept the conical bow shock at off-design Mach numbers. The system in principle obviates wave drag to provide shock-free supersonic flight with improved efficiency and no sonic boom.

CROSS-REFERENCE TO RELATED APPLICATIONS

This is a continuation-in-part of U.S. patent application Ser. No.07/557,418, filed Jul. 23, 1990, now abandoned.

BACKGROUND OF THE INVENTION

My previous U.S. Pat. No. 4,483,497, issued Nov. 20, 1984, disclosed amanifold arranged to emit a planar jet at a velocity greater than thatof flight and in the aft direction below and apart from a lifting wing.The energized jet fluid, supplied by the aircraft propulsion system,would be exhaust or bleed air from a turbojet/turbofan, a planar ramjet,or any other engine compatible with the aircraft propulsionrequirements.

The gap between the wing underside and this jet constitutes a planarnozzle which at supersonic speed decreases the flow velocity with acorresponding increase in pressure. The upper jet interface generates aseries of negative (counterclockwise) vortices whose strength isdetermined by the jet/gap velocity ratio. The lower jet interfacegenerates a series of positive (clockwise) vortices whose strength isdetermined by the jet/free-stream velocity ratio. Since the gap flowvelocity is less than that of the free stream, the vortices on the upperjet interface are stronger, and their excess strength can provide therequired negative circulation reaction to the positive circulationdeveloped by the lifting wing.

Effectiveness of the circulation reaction system provided by thisunderwing jet can be maximized by enhancing transformation of theresidual energy of the jet, which appears as linear velocity, into arotational form, i.e., by transforming the linear momentum into angularmomentum. This transformed jet then provides a more effective reflectionboundary to recover energy of the underwing compression waves. Fuselagebow compression energy is also recovered by reflection employing a nosering, possibly again in conjunction with a jet. This energy recoveryimprovement, based on reflection, is the subject of the presentinvention.

BRIEF SUMMARY OF THE INVENTION

The present invention outlines the recovery of supersonic compressionwave energy, generated both by the lifting wing and the fuselage bow, byits reflection onto a suitably inclined structural surface to providethrust and hence useful work.

For the wing, the invention provides means to selectively enhance mixingof the planar jet upper surface with the underwing gap flow to maximizethe negative angular momentum or circulation reaction, preferably withinthe chord length of the wing. These means include but are not limited toacoustic, electrical, mechanical, laser, and geometric systems to induceresonance in the eddies formed on the upper jet interface and therebyaccelerate the mixing process and vortex growth.

Such enhancement, by fully utilizing the available residual linearenergy in the jet, will minimize both the mass and velocity of the jetfluid required from the engines, thereby providing the required negativereactive circulation with the least demands on the propulsion system.Further, reduced jet mass flow will minimize the size, weight, and dragof the underwing manifold itself.

Enhanced growth of the upper interface eddies by mixing will alsoprovide the circulation reaction within the chord length of the wing,where the greater vortex growth rate presents a more favorableasymmetric boundary condition to increase the reflection of theunderwing compression waves back upwards to maximize the pressure on theupwards reflexed wing backside to provide lift and thrust. This mixingat the same time spreads the momentum over a greater mass of air therebyserving as a form of jet augmentation, where the wing backside serves asthe jet augmentation shroud.

Finally, the increased rotational momentum provided by this enhancedmixing augments the induced aft velocity of this upper vortex array asan expansion in the flow below, serving to cancel the forward velocityperturbation related to the compression imposed on the flow below thejet/vortex array due to its downward deflection. In fact, the jet/vortexarray formed and discharged aft of the wing has essentially zerovelocity with respect to the outer flow, and hence cannot form anywaves.

Enhanced mixing growth is also provided through inclining of the jetupwards toward the wing to increase reflection of underwing compressionwaves by forcing a turn by the gap flow through a greater angle withcorresponding increase in the pressure on the wing under surface.Mounting of the nozzle for pivotal motion to increase and decrease theangle of inclination at a resonant frequency is employed for furtherexcitation of the jet interface with the gap flow further enhancingshock wave reflection.

Creation of the jet with a velocity profile having lower velocity at theupper interface of the jet with consequent higher pressure and highervelocity at the lower interface of the jet with associated lowerpressure enhances underwing mixing on the upper surface of the jet whileproviding greater inertial on the lower interface of the jet to minimizeadverse defection of the jet downward.

For the fuselage, the invention provides a nose ring to intercept andreflect the conical axisymmetric fuselage bow shock back inwards onto asuitably inclined fuselage shoulder to recover its compression energyinto thrust and useful work.

These considerations thereby avoid formation and propagation of shockwaves to the ground to cause a sonic boom.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing and other readily apparent features of the presentinvention will be better understood by reference to the following moredetailed specification and accompanying drawings wherein:

FIGS. 1a and 1b are schematic representations of the vortex patternsgenerated by both subsonic and supersonic wings, contrasting thecirculation reaction requirements in these two flight regimes.

FIG. 2 is an illustration of the flow field generated by an examplesupersonic wing, delineating its shock and expansion waves and thevortex sheets produced by their intersections.

FIGS. 3a and 3b are a presentation of the wake characteristicscalculated for the example wing of FIG. 2, illustrating the velocitydefects imposed on the flow by the shock/expansion wave system, and FIG.3c is a tabulation of resulting wake circulation reactions forcomparison with the positive wing circulation developed as the wingadvances a unit distance.

FIG. 4a is a schematic presentation of both the present shock-generatingsupersonic system and the new shock-free system of the presentinvention, illustrating the turn reactions provided by the two systems,and FIG. 4b displays Crocco's Theorem which offers a mathematicalexplanation of how these systems differ.

FIG. 5 is a schematic illustration of the wing/jet system of the presentinvention, showing the wing, underwing jet manifold, and high velocityaft jet with its upper and lower interface vortex arrays, where theupper jet interface vortex array displays enhanced flow mixing and wavereflection.

FIG. 6 is an illustration comparing the boundary conditions provided bythe classical vortex sheet as presented in textbooks with the real fluidshear layer of the present invention.

FIGS. 7a and 7b are illustrations comparing reflection of supersonicwaves from a conventional symmetric jet with that of an asymmetric jetprovided by the action of a real fluid employing acoustic excitation.

FIG. 8 is a chart illustrating the dependence of reflected supersonicwave angles on the inclination of the reflecting surface.

FIG. 9 is a chart illustrating the increased compression wave reflectioncalculated for the enhanced mixing of the present invention.

FIG. 10 is a chart showing the deformation and compression of a vortexadjacent to a solid wall, and illustrating how the wall may berepresented by an image vortex.

FIG. 11 is a schematic illustration of how a deformed circular nozzlemay be represented as a transformation of a linear supersonic nozzle.

FIG. 12 is a summary chart illustrating the velocity and pressuredistributions of the upper and lower regions of the deformed vortex.

FIG. 13 is a mathematical representation for underwing supersoniccompression wave reflection as the sum of three terms, together withgraphical representations, showing the magnitudes of these terms as afunction of operating parameters.

FIGS. 14a, b, c, d and e are schematic presentations of severalelectromechanical excitation mechanisms for the section of the underwinggenerator shown by lines 14--14 in FIG. 5 to induce resonance in theinterface vortex array to enhance flow mixing and stimulate vortexgrowth.

FIGS. 15a, b, c, d, e, f, g and h are schematic illustrations of thesupersonic vortex growth mechanism, showing how the vortices, excited bya pressure signal, rotate, pair, and merge to enhance growth. FIG. 15aalso shows a receiver to measure the emerging frequency of the vorticesand transmit this data to control the excitation frequency.

FIGS. 16a and 16b are comparisons illustrating the two systems,contrasting the negative wake of the conventional dissipativeshock-generating system with the positive wake of the isentropicshock-free system of the present invention.

FIG. 17a is a schematic representation of the flow field aft and belowthe wing/jet system of the present invention, illustrating the aft wash(expansion) induced by the dominant upper vortex array to cancel theforward wash (compression) associated with the wave system of theconcave jet underside, providing a rotation-free downward momentum belowthe jet system generating lift with no sonic boom. FIG. 17b is agraphical representation of the ground pressure for the schematic ofFIG. 17a.

FIG. 18a is a chart illustrating the shock-free mechanism of the presentinvention in transforming the downward momentum function of a supersonicwing into a subsonic vortex array. FIG. 18b is a graphicalrepresentation of the ground pressure for the schematic 18a. FIG. 18c isan illustration of the mathematical transformation of the reactionsystem.

FIG. 19 is a profile view of an aircraft incorporating the underwing jetwith enhancement provisions of the present invention, and alsoillustrating forward wave containment and aft jet mixing arrangements tominimize fuselage wave drag.

FIGS. 19a and 19b are section views along lines 19A and 19B,respectively, in FIGS. 19a and 19b.

FIG. 20 is a schematic representation of a axisymmetric fuselage with abow shock reflection ring, a shock wave control nose spike foroff-design conditions, an energy recovery/thrust shoulder extendingcompletely around the fuselage, and a jet flap discharging aft from thereflection ring.

FIG. 21 is a perspective view of a corporate jet size aircraftillustrating both the forward-fuselage ring, bow energy recovery systemas well as the large-span, modest-swept wing with its underwing jet forwing compression energy recovery in accordance with the invention.

FIGS. 21a and 21b are section views along lines 21A and 21B,respectively, in FIG. 21. FIG. 21c is a front view of the embodiment ofFIG. 21.

FIGS. 22, 22a, 22b and 22c are section and front perspective views of atransport size aircraft illustrating the same features as in FIG. 21.

FIG. 23a is a depiction of conical bow shock with a retracted nosespike.

FIG. 23b is a depiction of the bow shock condition with the nose spikeextended.

FIG. 23c is a chart illustrating use of the extendable nose spike tocontrol interception of the conical bow shock wave by a forward fuselagebow ring at off-design conditions.

FIGS. 24, 24a, 24b and 24c are the perspective section and front views,respectively, of a transport size aircraft illustrating a parasol wingwith engines located at the focus of its inward inclined compressionwaves so their exhaust can reflect these waves back to the upwardreflexed backside of the wing for recovery into thrust.

FIG. 25a is a section view demonstrating the angle of inclination of thejet with respect to free stream.

FIG. 25b is a section view of the wing and nozzle with the nozzlepivoted to provide a higher angle of inclination in the jet.

FIG. 26a discloses a double nozzle configuration with a common singlestagnation stream to provide a negative pressure profile by modificationof the stream int eh nozzles.

FIG. 26b is a second embodiment of the double nozzle arrangementproviding two separate jets originating from two separate stagnationpressure plenum chambers.

FIG. 26c discloses a single unsymmetrical nozzle arrangement forcreating the desired negative velocity profile.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

In the following paragraphs like numbers refer to the same or similaritems from figure to figure.

The basic circulation reaction requirements imposed on the flow by alifting wing to satisfy conservation of angular momentum are illustratedin the schematic vortex patterns of FIG. for both subsonic andsupersonic speeds. The differing reaction systems shown follow from therestrictions on pressure transmission due to the speed of sound.

At subsonic speed the wing 20 initially turns a horizontal flowdownwards to develop lift via a positive circulation +Γ₁. Conservationof angular momentum then requires an equal and opposite sign reactivecirculation -Γ₁ which is represented as the so-called starting vortex-20r, with these two spanwise circulations connected by wing-tiptrailing vortices 22 to produce a closed vortex ring. Thereafter, as thewing 20 moves forward, it is preceded by its own pressure field,generating a forward upwash, which the wing turns downward to developthe lift force. For a two-dimensional wing the upflow forward of thewing and the downflow aft of the wing are symmetric, so there is zeronet downwash, the lift being produced by the change in linear momentumfrom the forward upflow to the aft downflow.

Since the pressure field is thereby carried along with the wing as thesame bound vortex +Γ₁, no further new circulation is developed, nor isany further reactive circulation required. The vortex ring, orrectangle, is thereafter simply stretched as the wing advances topositions 20b and 20c, leaving the initial starting vortex -Γ₁ farbehind in the fluid.

At supersonic speed the pressure field generated by the liftingcirculation +Γ₁ cannot precede the wing 23, which accordingly mustcontinually turn the flow down from the horizontal to generate newpositive circulations +Γ₂ and +Γ₃ as the wing advances. Conservation ofangular momentum then requires sustained negative circulation reactions-Γ₂ and -Γ₃. Thus, the positive wing lift circulation 23 is reacted bynegative circulation -23r, the positive wing lift circulation 24 isreacted by negative reactive circulation -24r, the positive wing liftcirculation 25 is reacted by negative reactive circulation -25r, and soon. Again the spanwise lift and reactive vortices are connected bywing-tip trailing vortices, but now generating a series of new vortexrectangles--a new rectangle for each unit distance of advance, ratherthan simply stretching the original vortex rectangle as in the subsoniccase. Of course, the wing advance, its circulation development, and thevortex reactions are all smooth and continuous; the unit advancepatterns cited are merely illustrative.

In FIG. 2 a flat plate wing 26 developing a positive lift circulation+Γ_(w) is illustrated generating shock waves 27l and 27u from the lowerand upper surfaces of the wing respectively. Expansion fans 28l and 28uare also shown emanating from the wing lower and upper surfacesrespectively.

Below the wing, intersections of the elements 28l(1), 28l(2), and 28l(3)of the lower expansion fan 28l with the lower shock wave 27l generatevortex sheets 29l(1), 29l(2), and 29l(3) respectively, whose sustainedvortex arrays all have a negative (counterclockwise) sense of rotation.

Above the wing, intersections of the elements 28u(1), 28u(2), and 28u(3)of the upper expansion fan 28u with the upper shock wave 27u similarlygenerate sustained vortex sheets 29u(1), 29u(2), and 29u(3)respectively, but now the sense of rotation of these vortex arrays ispositive (clockwise).

The slipstream 30 between the lower and upper flows aft of the wingtrailing edge 31, due to the higher velocity in the lower flow 32l, isalso shown in FIG. 2 generating a further negative reactive vortex array33.

The wake characteristics of the shock/expansion wave system of FIG. 2are illustrated in FIG. 3. The shock waves 27l and 27u are of uniformstrength until intercepted by the expansion fans 28l and 28u, degradingkinetic energy of the flow into heat, i.e., causing uniform velocitydefects 34l and 34u and temperature rises 35l and 35u. When theexpansion fans intercept the shock waves, the velocity defectsprogressively decrease, providing velocity variations, the integrals ofwhich are the reactive circulations.

Thus below the wing 26, after interception of the shock wave 27l by theexpansion fan 28l, the velocity defect 34l decreases progressively inthe negative y direction, providing a negative slope of the velocityprofile, the integral of which is the required negative circulationreaction -Γ_(l). Surprisingly, for the Mach 3 wing at 10 degree angle ofattack example case considered, calculations show that this negativecirculation reaction -Γ_(l) is almost five times that required tobalance the positive wing lift circulation +Γ_(w), as indicated by theircirculation areas.

However, above the wing 26, after interception of the shock wave 27u bythe expansion fan 28u, the velocity defect 34u decreases progressivelyin the positive y direction, providing a positive slope of the velocityprofile, the integral of which is a large positive circulation reaction+Γ_(u), almost four times that of the wing circulation +Γ_(w), again asindicated by their areas. This result explains the even larger negativecirculation reaction -Γ_(l) generated by the lower flow; it is requiredto balance this adverse upper flow reaction +Γ_(u) as well as thepositive wing circulation +Γ_(w). This system for the example casegenerates eight times the reactions required, including both thepositive and negative circulations, and hence dissipates eight times theenergy necessary to balance the positive wing circulation. The result isthe abnormally high wave drag of the shock-generating reaction system,and indicates that supersonic lift should be generated by the wing lowersurface but not its upper surface.

The calculated value 33 of the slipstream negative circulation reaction-Γ_(s) is also shown; it is small, only about one percent of the wingcirculation +Γ_(w).

Circulation reactions to the positive circulation +Γ_(w) developed by alifting wing, at both subsonic and supersonic speed, are provided byfree vortices in the flow. At subsonic speed the one-time startingvortex is generated by the wing trailing edge 31. At supersonic speed,as illustrated in FIGS. 2 and 3, and summarized in FIG. 4, a wing in auniform energy flow 35 employs intersections of the dissipative shockwaves 27l and 27u with the expansion fans 28l and 28u to generatecontinuous free vortex arrays, both positive and negative as shownabove, providing the net sustained negative circulation reactionrequired to balance the continuous generation of positive wingcirculation. The new variable energy circulation reaction system of thepresent invention employs residual energy from the propulsion system toprovide a higher velocity jet stream 36 to generate this requiredsustained free vortex array 37 on the interface between the free stream35 and the jet stream 36.

The distinction between these two circulation reaction systems isillustrated by combining the conservation equations of mass, momentum,and energy into a single equation known as Crocco's Theorem 38, where

T=temperature

s=entropy

q=velocity

h=enthalpy

This three-term expression illustrates the two and only two possiblesupersonic circulation reaction systems. For the two-dimensional wingconsidered, the second term 38b of this equation represents the verticalgradient of the rotational energy generated by a lifting wing in turningthe flow downward. In a uniform energy flow 35 the third term 38c,representing the vertical gradient of the constant total enthalpy, iszero, and hence the circulation reaction, as demonstrated by theresulting two term equation, must be provided by the first term 38a,which represents the energy dissipation of the shock wave system 27.However, inclusion of the energized underwing jet 36 provides anon-uniform vertical enthalpy gradient and hence a non-zero right-handterm 38c, which can balance the equation without a contribution from theshock-dissipative first term 38a, thereby achieving a shock-freecirculation reaction.

This shock-free circulation reaction system of the present invention isdisplayed in FIG. 5, illustrating a cross section view of a shaped wing39 having a flat upper surface 39u and a lower surface gently curvedconcave down in both its forward 39f and back 39b sections, connected bya convex center portion 39c, in all comprising the upper section of asupersonic nozzle. A manifold 40 is mounted below the wing emitting ahigh velocity planar jet 36 in the aft direction, generating vortexarrays 37l and 37u on its lower and upper jet interfaces respectively.Vortex growth provisions 41 are incorporated in the upper surface of themanifold 40 to selectively enhance mixing of the upper jet interfaceeddies 37u of the jet 36 with the gap flow 42, providing the largerupper vortex array 37u shown.

The high velocity planar jet 36 is comprised of energized fluid providedby the propulsion system, such as exhaust or engine bleed air. Thisfluid may be ducted spanwise from the engines through the wing leadingedge, providing thermal deicing, and/or in the forward plenum section40f of the underwing manifold 40.

The enhanced growth of the upper interface vortex array 37u is acritical feature of the present invention. First, this vortex array isendowed with increased strength by its generation within the wing chordlength where the slower underwing gap flow provides a greater velocitydifferential. The reduced velocity of the gap flow 42 is the driver forthis mechanism, and it is available only within the wing chord length.Aft of the wing the jet is bounded by the free stream on both itsinterfaces, and hence the vortex strength is the same on both sides,i.e., the increase in circulation reaction strength on its upper sidewill vanish. Second, its enhanced mixing with the gap flow transforms agreater portion of the residual linear jet energy into rotational form.These two provisions thereby make more effective use of the availableresidual jet energy and generate more of the required negativecirculation reaction than would otherwise occur.

The benefits of this enhanced vortex structure appear in several ways.First, it provides an improved boundary condition for the underwingflow. The enlarged upper vortex array 37u serves as a more effectivepneumatic shield in reflecting a greater portion of the compressionwaves 43 generated by the forward concave-down wing underside 39f backupwards to increase the pressure on the wing backside 39b, providing athrust component due to its upward reflex. From another point of view,the gap flow 42 is restricted by the vortex growth to a channel oflesser cross section, generating a higher pressure which is sustained tothe wing trailing edge 31. Second, the enhanced mixing spreads theresidual jet energy over a greater mass of air, providing a form of jetaugmentation to increase the propulsive efficiency of the underwing jet,with the upwards reflexed wing backside 39b acting as the jetaugmentation shroud. Third, the underwing jet/vortex array extends faraft of the wing providing a longer effective wing chord to improve liftefficiency.

The vortex growth increases the wave reflection 44 by providing adiverging upper boundary 37b having improved reflection over theclassical boundary. Wave reflection from a free surface or vortex sheetis normally represented by the classical boundary conditions whichapproximate the boundary as having infinitesimal thickness, with thepressures 50c and flow directions 51c the same on both sides of thesheet, as shown in the upper portion of FIG. 6. However, with enhancedmixing the vortices or eddies grow rapidly in the flow direction tofinite size which can sustain a transverse pressure gradient 50rf byvortex deformation and produce flows on its two sides that diverge 51rfin direction, as shown in the lower portion of FIG. 6. These real fluidboundary conditions increase the reflection and hence recovery of thewing compression energy, compatible with the increased negativecirculation reaction of the larger size vortices.

The underwing jet is shown in FIG. 7 for these two boundary conditionsincluding example Wave patterns illustrating the greater net reflectionfrom the asymmetric jet:

Case 1 - Classical Boundary Conditions - Symmetric Jet

Case 2 - Real Fluid Boundary Conditions - Asymmetric Jet

For Case 1 an example is presented for a reasonable jet velocity ofabout twice the flight velocity. A compression wave 43 generated by theforward wing undersurface 39f encounters the upper surface 37u of theunderwing jet 36. One third of the wave is reflected upwards from thisclassical symmetric jet 36c as a compression wave 44cr; two thirds istransmitted downwards into the jet as a compression wave 44ct at alesser inclination due to its higher Mach number. On encountering thelower surface 37l of the underwing jet 36, which is a weaker interfacedue to the lesser difference in velocities across its surface, only onefifth of the wave is reflected and this time as an expansion wave 44cre(shown dashed). This wave continues upwards, encountering again thestronger upper interface 37u, where as before one third is reflected andtwo thirds transmitted. The final upwards transmitted wave is anexpansion having a strength which is the product of its precursors, orabout one tenth, which detracts from the one-third initial compressionreflection.

For Case 2 the same mechanism holds. But the real fluid boundaryconditions of the asymmetric jet, for the same example where the jet hasa velocity twice the flight velocity, provide an initial reflection oftwo thirds and a subsequent upwards transmitted expansion of one fortyfifth, almost negligible. So the real fluid asymmetric jet 36rf gains intwo ways, first by a greater initial reflection and second by a lessersubsequent expansion.

The mechanism of flow divergence due to mixing layer growth andinclination of the real fluid boundary is shown in FIG. 8. An incidentMach wave 54 at angle μ and an incident weak compression wave 56 atangle β are shown. The flow U turns through an angle Θ₁ on passingthrough the weak compression wave. The angles of the outgoing reflectedwaves are shown to depend on the inclination of the reflecting surface.For a surface inclined at the same angle Θ₁ as the flow U the wave willnot reflect at all and instead be "absorbed", or sometimes is said toreflect as a Mach wave β₁. If the surface is not inclined, shown as Θ₂,the weak compression wave will reflect at an angle β₂ similar to theangle of incidence β. A surface inclined upwards, say at an angle Θ₃ asshown, provided by a real fluid diverging asymmetric jet, will reflectat an increased angle β₃ and turn or deflect the flow U upwards at acorresponding angle Θ₃, with a proportional increase in pressure. Thispressure increase will also deflect the underwing vortex flap downwardsto a greater extent, thereby increasing the downward momentum impartedto the flow below.

This increase in pressure due to vortex spreading is shown in the chartof FIG. 9, which is a plot of the reflection coefficient R versus thedynamic pressure ratio λ of the two streams. In this plot, whichconsiders only the flow spreading real fluid boundary condition 51rf,the parameter κ is the ratio of the vortex spreading angle Θ to thedownward angle σ of the forward wing undersurface 39f. The plot showsthat for a unit value of this ratio κ, which corresponds to a vortexspreading angle equal to the undersurface slope, and for a dynamicpressure ratio λ of 1.85, the reflection is increased by a factor ofthree over the value for the classical boundary.

The increase in reflection due the ability of a real fluid shear layer36rf to sustain a pressure gradient depends on vortex deformation, asillustrated in FIG. 10. In this figure, taken from a standard textbook,a vortex 80 is shown in the proximity of a solid wall 82, where thevortex is deformed to a greater extent in its side 80a adjacent thewall. This greater deformation reduces the radii of curvature 84 of therotating flow in the compressed region, increasing the centrifugalforce, which is balanced there by a greater pressure gradient. FIG. 10also illustrates that the solid wall 82 can be replaced by an imagevortex 86 to facilitate calculation of the flow field.

The pressure gradient mechanism in a deformed vortex 80 is furtherillustrated in FIG. 11, which displays its analogy with a conventionalsupersonic nozzle. In this figure a linear supersonic nozzle 82 is shownwhere the flow is compressed in its throat O-A, providing an increase inpressure with a corresponding decrease in velocity. If this nozzle 82were wrapped around its center 0 into a circular nozzle 80 as indicated,the transformation would represent the deformed circular nozzle 80described.

Vortex deformation effects are illustrated in FIG. 12. The upper sketchshows the deformed vortex 80 together with a plot of its verticalvariations in velocity 90 and pressure 92. The middle plot illustratesthe velocity distributions of the upper 80u and lower 80l regions of thevortex 80. The lower plot shows the pressure distributions of the tworegions 80u and 80l in a similar manner.

The two further reflection factors imposed by the real fluid boundaryconditions, vortex spreading and vortex deformation, may be examined ina quantitative manner by deriving a mathematical representation, basedon substitution of the flow potential into these new boundary conditionsand solving for the resulting reflection.

Such a derivation results in a new General Reflection Coefficient R asshown in FIG. 13, presenting reflection from the real fluid shear layeras a function of the geometric and flow variables, expressing thisreflection as the sum of three terms, namely:

Classical 102 - the effect of an idealized infinitesimal thickness layer

Pressure Gradient 104 - the effect of the deformed real fluid vortices

Spreading 106 - diverging surface effect due to Vortex growth

In addition to the classical term 102, two new terms, "pressure gradient104" and "spreading 106" now appear, one for each real fluid boundarycondition, pressure 50rf and flow direction 51rf. The factors in thisequation are defined as follows:

    ______________________________________                                        Symbol       Factor                                                           ______________________________________                                        λ     dynamic pressure ratio                                           α      total spreading angle                                            σ      underwing down-angle                                             δ      vertical vortex pressure gradient,                               p.sub.gap    p.sub.∞                                                    β       (M.sup.2 - 1).sup.1/2                                            M.sub.∞                                                                              free stream Mach number                                          M.sub.jet    jet Mach number                                                  M.sub.gap    gap Mach number                                                  Ε    vortex deformation, %                                            p.sub.gap    underwing gap pressure                                           p.sub.∞                                                                              free stream pressure                                             ρ        air density                                                      ______________________________________                                    

The spreading, or vortex growth term 106, depends only on the angles αand σ, and the dynamic pressure ratio λ. The underwing down angle σ isnegative, so the spreading pressure contribution is positive. Itsincrease in pressure is achieved primarily by deflection of the flow dueto the upwards inclined boundary, as shown in FIG. 8 previously.

The pressure gradient term 104 depends primarily on the pressuredifference δ between the underwing gap flow M_(gap) and the externalfree stream M.sub.∞, which is the vertical pressure gradient p.sub.∞-p_(gap) across the shear layer leading to vortex deformation, asillustrated previously in FIG. 12. This pressure gradient will besustained by the internal vortex structure, which balances thecentrifugal force within the deformed vortex by the pressure gradient.In a precise model the gradient will appear in the stream direction orinclined plane of the underwing vortex flap, with both vertical andhorizontal components. The horizontal component will transport thepressure aft, providing a longer effective wing chord.

The plots in the lower portion of FIG. 13 illustrate the dependence ofthe reflection coefficient R on its three terms. The classical term 102is presented in standard textbooks, and provides only a modestreflection. The pressure gradient 104 and spreading 106 terms are thecontributions of the real fluid boundary conditions provided by thepresent invention. These contributions appear considerably larger thanthe classical term 102 for reasonable values of the parameters shown,illustrating the major gains in performance achieved by the newshock-free supersonic system of this invention.

The enhancement provisions of the present invention are based onimparting a pressure pulse into the flow at a frequency related to thenatural frequency of the eddy formation on the jet interfaces. Theexcitation frequency may be the same as the natural frequency, or someharmonic thereof. The excitation provisions may be acoustic, electrical,mechanical, geometric, laser, or any other oscillation means such as toemit a pressure perturbation into the flow in a manner as to induceresonance in the interface eddies, providing an unstable internalstructure to cause the eddies to merge, pair, and grow rapidly in thestreamwise direction. FIG. 14 illustrates several possibleelectromechanical mechanisms for imparting this excitation pulse.

FIG. 14A illustrates a pulse mechanism 140 mounted in the top wall ofthe manifold throat 142 to emit a pressure pulse downwards from aplunger into the manifold internal flow; FIG. 14C has a similar plunger140 internally mounted to pulse a flexible membrane 144. FIG. 14B showsa pulse mechanism 140 similarly located but oriented to emit the pulseaft. FIG. 14D illustrates an excitation mechanism 140 to emit a pulseinto the gap flow 42 above the manifold. FIG. 14E illustrates a steppedupper inside surface 148 in the nozzle to impart vorticity by localseparation selectively in the upper surface shear layer.

Acoustic excitation mechanisms could also be provided in the manifold insimilar locations, primarily in the upper section of the planar nozzle40 to selectively excite the jet stream 36 on its upper interface 37u.

A notch 146 in the upper wall of the manifold 40 would also serve as anexcitation mechanism, and would have the advantage of being a passivestructure, operating like a whistle. Variable width of the notch wouldbe incorporated to control its frequency.

The excitation mechanisms could also be located on the forward underwingsurface 39f, ahead of the compression field, so that its excitationpulses would interact with the upper jet interface in conjunction withthe compression waves 43 to enhance the vortex growth.

FIG. 15 illustrates schematically the vortex pairing character providedby the excitation mechanism disclosed. In this sketch the vorticesformed on the upper jet interface 37u between the underwing gap flow 42and the jet flow 36 are numbered in sequence as they emerge from thenozzle 40. The excitation mechanism 41 then emits pressure pulses at thesame rate, i.e., the same frequency or some harmonic thereof, to depressevery other vortex, alternating between overpressure and underpressure,causing the vortex pair to rotate. As succeeding vortex pairs emerge,the continuing vortex rotation leads to their pairing and providing astreamwise growth structure. From another point of view the excitationprovisions within the nozzle 40 will cause the jet 36 to emerge with analternating underpressure and overpressure, which with proper harmonicsequencing, will generate the vortex rotation and pairing mechanismdescribed.

The receiver 54 could be located downstream of the jet nozzle, say onthe wing undersurface, as shown in FIG. 5, to detect the emergingfrequency of the interface vortices, and transmit this data to controlthe frequency of the exciter 41.

The two circulation reaction systems--shock-generating and enhancedshock-free--are compared in FIG. 16. The shock-generating system of FIG.16A has a negative wake 34n; the shock-free system of FIG. 16B has apositive wake 34p. The shock-generating system produces opposingcirculation reactions, both positive and negative, operating at crosspurposes, involving energy dissipations several times that required toproduce a net negative circulation reaction to balance the positive wingcirculation +Γ_(w). This energy drain, since it is extracted from theflow at the price of wave drag, is particularly severe on performance.

The shock-free system, on the other hand, has a positive wake 34p, whichnevertheless has a negative slope of its velocity profile, the integralof which is the required negative circulation reaction -Γ_(r). Further,by employing a flat upper wing surface 39u, lift generation that wouldproduce an adverse reaction is avoided. The system thereby employsminimum energy to generate only the negative circulation reaction -Γ_(l)required in the lower wing gap flow, aside from a small positivereaction +Γ_(b) due to the jet bottom interface. Furthermore, thenegative circulation reaction provided by this positive wake isgenerated by residual energy from the propulsion system, with no adverseeffect on performance.

The improved shock-free system of the present invention is directedtowards employing enhancement provisions to force the transformation ofresidual linear jet energy into rotational form completely within thewing chord length. Such provision of the complete negative circulationreaction results in recovery of the underwing compression energy intoincreased pressure on the wing backside, as shown in the paragraphsabove.

The jet/vortex array assembly will extend aft far beyond the wing asshown in FIG. 17, and will be deflected downwards by the underwingpressure while within the Mach wedge extending from the wing trailingedge. Thereafter this assembly will be gradually deflected back towardsthe horizontal by the free stream below, thereby providing a much longereffective wing chord to spread the downward momentum over an increasedmass of air to improve the lift efficiency.

In the forward region exposed to the underwing lift pressure, downwarddeflection of the vortex assembly possibly would generate compressionwaves 150 having a forward perturbation velocity (compression) 152 aslong as the assembly is within the wing Mach wedge, as shown in FIG. 17.These compression waves would normally coalesce into a strong shock waveextending to the ground and cause a sonic boom. However, with thepositive wing circulation completely balanced by the negative jet vortexcirculation reaction, further negative circulation such as would beproduced by shock waves cannot arise in the flow. Hence generation offurther shock waves in the flow below this assembly is not possible.

The answer to this paradox is illustrated in this figure, showing thatthe induced aft velocity (expansion) 154 of the dominant upper interfacevortices 37u cancels the forward perturbation velocity 152 associatedwith the compression waves 150, generating a rotation-free downwardmomentum to provide lift but not allowing the compression waves tocoalesce into a shock wave. Hence there is no shock wave extending tothe ground to cause a sonic boom. Instead the weight of the aircraftappears on the ground as a long, smooth, low pressure footprint 156extending far aft of the wing.

FIG. 18 illustrates another point of view regarding the jet/vortex flapassembly mechanism in the region aft and below the wing. As shown inthis figure, the vortex flap assembly 37u is discharged aft from thewing with a velocity increment to provide thrust, but which aftersubstantial mixing as shown has only a modest residual velocity. Thusafter discharge the flap has essentially no horizontal motion, i.e., itis simply laid down as the wing passes. If it were attached to the wing,it would have a high velocity relative to the outside stream. But it isnot attached to the wing, and has essentially zero velocity relative tothe outside air into which it is discharged. Hence this flap willgenerate no waves.

In a sense the mechanism thus transforms a supersonic wing into asubsonic vortex array having an effective Mach number 160 of zero. Thevortex flap 37u will be deflected downwards by the underwing pressure,and will in turn push down the outer fluid. So linear downward momentumis imparted to the air below, providing lift, but no waves are generatedassociated with rotational momentum.

Again this model spreads the downward momentum over a much longereffective wing chord 162, which in turn spreads the pressurecorresponding to the weight of the aircraft smoothly over a much longerfootprint 156 on the ground with no sonic boom.

A summary of the wing energy recovery system thus illustrates that thepresent invention provides three mechanisms in a synergistic manner thatincrease the pressure on the upwards reflexed wing backside to producethrust and useful work to benefit aircraft efficiency:

Underwing Compression Wave Reflection

recovers wave compression into useful work this recovered energy appearsas pressure on wing backside

Jet Augmentation

spreads jet linear momentum over increased mass of air

this mixing transforms linear momentum into rotational form

reduced mixed stream velocity augments pressure on wing backside

Jet Flap

increased underwing gap pressure

inclines shear layer downward

compresses and deforms shear layer vortices

deformed vortex internal structure sustains pressure gradient

this gradient has vertical and horizontal components

spreads lift momentum over longer effective wing chord

results in increased pressure on wing backside

Wave drag of the fuselage may also be alleviated by application of thereflection principles of the present invention. FIG. 19 illustrates acorporate jet aircraft having a cowl or jet manifold 170 mounted abovethe forward fuselage, which reflects the upwards oriented fuselage bowwave back downward to provide an increased pressure and thrust on theaft side of the bump region 172 on the upper fuselage surface, therebyrecovering into useful work the otherwise wasted bow wave energy.

The upswept aft underside of an aircraft fuselage is usually a region oflow pressure and hence high drag. The present invention locates afurther jet manifold 174 forward of this surface to emit a jet aft,employing excitation provisions as disclosed herein, to enhance vortexgrowth on the upper interface of this jet to provide a high pressure andhence a thrust on the lower fuselage backside 176.

The cowl of FIG. 19 may be extended completely around the fuselage, andmay include a manifold ring 180 to emit a jet flap 182 as illustrated inFIG. 20. This figure displays a circular fuselage 184 with a shoulder186 completely around its perimeter to recover the axisymmetric conicalbow shock 188 into forward thrust T and useful work. An extendable,forward, shock wave control spike 192 is also shown to locate the shock188 on the manifold ring 180 properly for reflection recovery atoff-design conditions.

A corporate size aircraft with such a bow shock reflection ring 180,nose spike 192, and inward inclined energy recovery shoulder 186 isshown in FIG. 21, with tail mounted engines 194t. The figure alsoillustrates that recovery of the wing shock wave energy as outlined inthis invention permits the use of a large-span modest-sweep wing 196,providing improved take-off and landing characteristics and less induceddrag at both subsonic and supersonic speeds. The bow shock ring 180 ismounted on a plurality of struts 198 such that the lower portion of thering 180l may be folded upwards for take-off and landing.

FIG. 22 illustrates the same principles of this invention applied to atransport size aircraft. The engines 194w in this case are mounted underthe wing.

Operation of the conical bow shock wave control spike 192 is shown inFIG. 23, presenting a Mach number Table showing for a conventionalwide-body transport fuselage the ring diameters required for compressionenergy recovery on the fuselage shoulder. This wave angle/ring diametergeometry is illustrated in FIG. 23(a). The extendable nose spikeaccordingly is arranged as shown in FIG. 23(b) to locate the conical bowwave on a 40.9 foot diameter ring over a flight range of Mach 2.4 downto Mach 1.8, thereby avoiding larger rings for lower Mach numbers wherethe entropy loss is small. The system then intercepts the conical bowshock wave and recovers its energy even at off-design Mach numbers.

The wing section 39, as shown in FIG. 5, has a flat upper surface 39uthat generates no lift at supersonic speed when operating parallel tothe free stream. However, the wing is provided with leading 200 andtrailing 202 edge flaps as shown in FIGS. 21 and 22, which do producelift on the upper surface when extended at low subsonic speeds,particularly for take-off and landing. The underwing jet 36 may also beemployed to increase lift during take-off by blowing the extendedtrailing edge flaps 202. Further, this dispersion of the propulsive jetwill decrease noise at take-off

The spanwise underwing jet disclosed in this application may becontinuous or intermittent. The aircraft illustrated in FIGS. 21 and 22display a continuous spanwise manifold emitting a jet across essentiallythe entire span of the wing and continuing uninterrupted under thefuselage.

The underwing jet alternatively may be provided in an intermittentmanner, say in the form of discrete segments, and in conjunction with awing shaped to focus the compression waves on those discrete jetsegments. FIG. 24 illustrates a transport size aircraft 210 having awing 212 concave down spanwise on both sides of the fuselage 214 toincline the wing compression waves 216 generated by the forward wingundersurface 212f inwards onto the underwing engines 218 and theirexhaust streams 220. The exhaust streams 220 then serve as thesediscrete jet segments to intercept and reflect the underwing compressionwaves 216 back to the upwards reflected wing backside 212b to recovertheir energy into useful work.

Use of the entire exhaust for this reflection role with its large energymay preclude the necessity of interface mixing as outlined in thisapplication, but if required the same enhancement provisions may beapplied to the upper surface of the exhaust jets as described herein.

Upward inclination of the jet with respect to the free stream adds tothe efficacy of the present invention. As shown in FIG. 25a the jetmanifold 40 emits the jet 36 at an average angle α with respect tot ehfree stream. In the embodiment shown in the drawings, this angle may beas much as 15°. The inclination of the jet forces the higher velocityflow from the manifold into the slower gap flow resulting in mixing withand entrainment of the slower gap fluid. A net upward inclination of thejet upper interface increases reflection of the underwing compressionwaves by forcing them to turn through a greater angle thereby requiringa greater turning of the gap flow as previously described with respectto FIGS. 5 and 7. Greater turning of the gap flow will result in acorresponding increase in pressure on the reflex undersurface of thewing.

As demonstrated in FIG. 25b, pivotal mounting of the manifold allowscontrolled alteration of the angle of inclination α. Varying of theinclination angle through rotation of the manifold at a resonantfrequency provides an alternate means of enhancement for the vortex flowat the interface of the jet. Downward loading on the jet manifold due tothe inclined jet is overcome through the enhanced reflection of theunderwing compression waves and increased pressure on the wingundersurface, which involves significantly greater mass of air toproduce increased thrust and lift more than off-setting the down-loadingon the jet manifold.

Enhanced stability of the jet and increased performance in reflection ofcompression waves to the wing undersurface is accomplished by creating avelocity profile having a lower velocity at the upper interface of thejet and a higher velocity at the lower interface of the jet. Thisnegative velocity profile locates the higher velocity portion of the jetat the lower interface where its greater inertia will minimize adversedeflection of the jet downward. The lower velocity upper surface of thejet with corresponding higher pressure will increase the beneficialexpansion of the jet upward to increase the desired pressure on the wingundersurface. FIG. 26a demonstrates a first nozzle system employing adouble nozzle to achieve the negative velocity profile. The nozzlesystem 230 comprises an upper nozzle 232 and a lower nozzle 234. Acommon stagnation pressure plenum chamber 236 provides high pressurefluid to be expelled through the nozzles. A retardation screen 238 ispositioned across the entrance to the upper nozzle to reduce thestagnation pressure of the gas stream provided from the plenum chamber.The upper jet is then emitted with a lower energy level and consequentlower velocity than the jet from the lower nozzle. The velocitydifferential is alternatively created through the use of a fuel burningsystem 240 at the entrance to the lower nozzle. The fuel burning systemadds energy to the steam entering the lower nozzle providing highervelocity flow in the lower nozzle to obtain the negative velocityprofile. Combination of the retardation screen and fuel burning in therespective nozzles further enhances the desired negative velocityprofile. The jet emitted by the two nozzles is demonstratedschematically by velocity profile 242 immediately aft of the nozzles.This discontinuous flow homogenizes to the desired velocity profile 244downstream of the nozzles. Further enhancement of the negative velocityprofile is accomplished through truncation of the upper nozzle to emitthe jet in an under expanded condition. The lower nozzle is nottruncated, fully expanding the flow at the jet lower surface for maximumvelocity.

FIG. 26b provides an alternate embodiment employing an upper and lowernozzle to obtain the desired negative velocity profile. Upper nozzle232' if fed by a stagnation pressure plenum chamber 246 while lowernozzle 234' is fed by a second stagnation pressure plenum chamber 248.Two separate streams of energized air of differing stagnation pressuresare drawn from the engine to feed plenum chambers 246 and 248. Employinghigher energy air in the lower plenum chamber provides higher velocityfor the fully expanded lower nozzle while use of lower energy air in theupper stagnation chamber provides lower velocity from the upper nozzle.A velocity profile 244' is achieved downstream of the nozzle.

A single nozzle arrangement to provide the desired negative velocityprofile is shown in FIG. 26c. The nozzle 250 incorporates a singleplenum chamber 252 for high energy air from the engine. Truncation ofthe upper expansion surface 254 of the nozzle in comparison with thefully expanded lower surface of the nozzle 256 provides the desirednegative velocity profile. Nozzle tailoring employing an asymmetricthroat or other fluid flow management techniques are alternativelyemployed in a single nozzle concept to provide the negative velocityprofile.

Having thus described my invention, what I claim as novel and desire tosecure by Letters of Patent of the United States is:
 1. In an aircrafthaving a wing, a nozzle, and a fuselage, a compression wave energycontrol device comprising:a forward bow introduced in said fuselage, andan aft tail, a bow ring mounted around at least a portion of saidforward fuselage bow to intercept and reflect a shock wave from saidfuselage bow aft onto a shoulder which is inwardly inclined toward acenterline of the fuselage on at least a portion of said fuselage whichreceives the reflected shock for pressure recovery into thrust anduseful work.
 2. The improvement claimed in claim 1 wherein the said bowring contains a nozzle, said nozzle emitting an energized jet aft andinwards providing a jet flap to contain the pressure of the reflectedshock.
 3. The improvement claimed in claim 1 wherein the said fuselagecontains an extendable nose spike to control the pattern of the bowshock wave generated by said fuselage for off-design conditions.
 4. Theimprovement claimed in claim 3 wherein said bow ring contains a sensorto control the extension of said nose spike.